The present invention relates to a gas turbine cooled blade and more particularly to an internal convectively cooled blade for a gas turbine. A leading edge cooling, particularly the cooling of an inner surface of a leading edge of the turbine blade, of a multipass convectively cooled gas turbine blade can be improved.
In a gas turbine, hot and high temperature combustion gas generated through a compressor and a combustor drives the gas turbine. With the improvement of engine thermal efficiency of the gas turbine in recent years, gas temperature now exceeds the heat resistance limit of the turbine hot components including turbine blades.
To cope with excessive thermal stress and corrosion generated in such a hot gas atmosphere, and to ensure sufficient mechanical strength and reliability, consistent improvements in heat-resistant alloys and coating materials and also in cooling technologies for the turbine hot components have been attempted.
Particularly because the turbine blade is directly exposed to the combustion gas and must have high cooling performance, an intricate cooling structure wherein the turbine blade is made hollow to form internal cooling passages and ejection holes to the outside and further cooling fluid such as a bleed from the compressor is introduced into both passages, has been employed.
In order to fully accomplish the high thermal efficiency by increasing gas temperature, an increase in total cooling flow rate that reduces the above efficiency significantly, especially an increase in the turbine blade cooling flow rate to which accounts for high percentage of cooling flow, must be minimized. Accordingly, various high performance gas turbine blades have been developed keeping pace with the improvement of blade fabrication techniques.
One of the typical gas turbine cooled blades is an impingement cooled blade disclosed in Japanese Patent Laid-Open No. 9623/1981. This impingement gas turbine cooled blade has a double structure consisting of a hollow blade main body and an insert body for cooling which is disposed in the cavity of the blade main body. The cooling fluid supplied into the cooling insert body flows through the jet ports of the surface and impinges against the inner surface of the turbine blade, thereby achieving high level of impingement cooling.
Though this impingement cooling structure of the gas turbine blade involves the disadvantage that the cooling insert body must be produced separately from the hollow blade main body, it has the advantage that the cooling performance for the turbine blade can be easily set in accordance with an external heat load distribution by adjusting the size and arrangement of the array of the jet ports, and it has high cooling performance for the gas turbine cooled blade on an average.
Another typical gas turbine cooled blade is a multipass convectively cooled blade disclosed, for example, in British Pat. No. 2112467 and United States Patent No. 4514144. In this gas turbine cooled blade, at least one structure of the cooling passage is defined inside the turbine blade in such a manner as to start from the root of the blade and to pass or extend and return in the direction of blade height. The cooling fluid flows along the inner surface of each blade wall, on which the projections and the pin fins for the heat transfer promotion are disposed, and cause increased forced convective cooling.
From the aspect of the blade fabrication techniques, this multipass convectively cooled structure of the gas turbine blades has the advantage that the blade main body and the most portions of the cooling passages can be molded all together by the precision casting, but it also involves the problem that adjustablity of the cooling performance in accordance with average cooling performance and the external heat load is lower than that of the impingement cooled blade.
Therefore, technical improvements inclusive of a new blade fabrication method have been attempted in order to improve the cooling performance for the gas turbine cooled blade by utilizing fully the heat transfer promotion elements or ribs and by miniaturizing the cooling passages.
Still another cooling structure of the gas turbine blades is a film cooling structure which injects a cooling fluid onto the blade surface to reduce the gas temperature on the external surface of the blade. However, this film cooling structure of the gas turbine blade is limited to apply to an airplane engine field or the like which uses a high quality oil free from plugging of the jet ports, and is therefore used conjointly with the cooling structure described above.
Since the following description is primarily directed to the cooling structure of the turbine bucket which can be adapted to low quality oil, the impingement cooled blade and the multipass convectively gas turbine cooled blade without the conjoint use of the film cooling structure will be examined.
In accordance with recent aerodynamic design of turbine blades, a blade outer shape having a smaller blade thickness than the conventional turbine blade has been employed in view of increasing aerodynamic performance, and the leading edge of the blade has an outer shape approximately to an ellipsis while the rear edge is made as thin as possible. In the moving blade, the aerodynamic design of each section is made in conformity with the difference of the flow conditions on the basis of the difference of a peripheral speed in the direction of blade height, so that the blade of the gas turbine is likely to have an outer shape which is twisted in the direction of blade height.
When a cooling design of the moving blade having the blade outer shape with excellent aerodynamic performance is made, cooling of the leading edge of the blade becomes particularly difficult if enough blade wall thickness is provided to ensure sufficient strength and reliability.
Fundamentally, this results from the fact that since the radius of curvature of the leading edge outer surface of the blade is relatively small, the external heat load increases and the heat transfer area ratio between the inner and the outer surfaces of the leading edge of the blade becomes small. It also exerts the following adverse influences upon each cooled blade.
Namely, in the case of the impingement gas turbine cooled blade, it is very difficult to design and mold the cooling insert body which is thinly twisted in conformity with the blade outer shape. Even if the blade outer shape is corrected at the sacrifice of the aerodynamic performance to some extent in order to have the cooling insert body insertive, the distance between the inner surface of the leading edge wall of the blade and the jet ports on the leading surface of the cooling insert body becomes too long and that causes reduction of the impingement cooling performance due to diffusion of the jet stream.
Therefore, from the aspects such as an aerodynamic performance, cooling performance and blade fabrication capability, the impingement cooled blade has both merit and demerit and the following multipass convectively cooled blade for the gas turbine is more generally used for practical application.
In the multipass convectively cooled blade for the gas turbine, the limitation of the aerodynamic design and the blade fabrication is not so severe, but the following problem occurs concerning the cooling design. Namely, inside the leading edge cooling passage having a roughly triangular section, the inner surface area of the leading edge of the blade reduces corresponding to the apex of an acute angle, the distribution of the cooling fluid to the inner surface of the leading edge wall decreases and the substantial flow velocity is comparatively low. Accordingly, the convective cooling performance of the multipass convectively cooled blade for the gas turbine cannot be improved even if the heat transfer promotion ribs are used.
The conventional multipass convectively cooled blade for the gas turbine described previously proposes also a method for improving the convective cooling by the heat transfer promotion ribs. British Pat. No. 2112467 uses slanting heat transfer promotion ribs on the rear side wall and body side wall towards the inner surface of the leading edge wall with respect to the flow and improves cooling of the inner surface of the leading edge wall by causing channelling. U.S. Pat. No. 4,514,144 defines spacings between the heat transfer promotion ribs disposed slantingly to the flow in order to promote heat transfer and reduce the fluid resistance.
However, the cooling improvement method for the turbine blades having these heat transfer promotion ribs primarily causes channelling and turbulence augmentation of the boundary layer of the cooling fluid in the proximity of the inner surface of the cooling fluid and promotes heat transfer, and its performance, particularly the performance when applied to the enhancement of cooling of the inner surface of the leading edge wall, is limited.
Accordingly, in the multipass convectively cooled blade for the gas turbine having less limitations from the aspects of the aerodynamic design and blade fabrication and being excellent in the overall evaluation, it is primarily the leading edge cooling passage that needs to be improved most essentially in order the further improve the cooling performance, particularly the cooling performance of the inner surface of its leading edge wall of the turbine blade.